Field of the Invention
The invention relates to a turbine nozzle in a turbine engine such as a turboprop or a turbojet.
Description of the Related Art
Conventionally, a turbine engine comprises from upstream to downstream a compressor feeding air to a combustion chamber, which ejects hot gas into a high-pressure turbine. The high-pressure turbine has at least one nozzle followed by a turbine wheel. The nozzle is sectorized and comprises inner and outer platforms between which there extend substantially radial airfoils. The nozzle serves to provide proper guidance of the hot gas leaving the combustion chamber towards the wheel.
In operation, it is observed that the temperature of the hot gas from the combustion chamber is not uniform circumferentially, thereby giving rise to hot points being formed on the inlet nozzle of the high-pressure turbine. In order to remedy that difficulty, it is known to form through orifices in the upstream neighborhood of the inner and outer platforms of the inlet nozzle, these orifices being distributed around the circumference of each platform and enabling air that has passed around the combustion chamber to be reintroduced into the high-pressure turbine. These streams of air serve to cool the upstream ends of the inner and outer platforms and to make the nozzle inlet temperature more uniform circumferentially.
Nevertheless, that type of technology is not satisfactory since the air passes through the inner and outer platforms at relatively high speed, which leads to the air penetrating radially a long way between the inner and outer platforms of the nozzle. As a result, this cooling air becomes mixed very quickly with the hot gas from the combustion chamber, and it can no longer provide good cooling of the inner and outer platforms.